U.S. patent application number 11/719911 was filed with the patent office on 2009-04-30 for inflatable bleed valve for a turbine engine.
Invention is credited to Brian Merry, Lawrence E. Portlock, Gabriel L. Suciu.
Application Number | 20090110544 11/719911 |
Document ID | / |
Family ID | 36829861 |
Filed Date | 2009-04-30 |
United States Patent
Application |
20090110544 |
Kind Code |
A1 |
Suciu; Gabriel L. ; et
al. |
April 30, 2009 |
INFLATABLE BLEED VALVE FOR A TURBINE ENGINE
Abstract
A compressor for a turbine engine includes an inflatable bleed
valve that selectively bleeds core airflow from the compressor. The
bleed valve has an inlet leading from the compressor and a
passageway leading from the inlet. An inflatable valve selectively
obstructs the passageway based upon a controlled supply of high
pressure air to the inflatable valve. The supply of high pressure
air may be compressed core airflow from an area downstream of the
inlet to the bleed valve.
Inventors: |
Suciu; Gabriel L.;
(Glastonbury, CT) ; Portlock; Lawrence E.;
(Bethany, CT) ; Merry; Brian; (Andover,
CT) |
Correspondence
Address: |
CARLSON, GASKEY & OLDS/PRATT & WHITNEY
400 WEST MAPLE ROAD, SUITE 350
BIRMINGHAM
MI
48009
US
|
Family ID: |
36829861 |
Appl. No.: |
11/719911 |
Filed: |
December 1, 2004 |
PCT Filed: |
December 1, 2004 |
PCT NO: |
PCT/US04/39989 |
371 Date: |
May 22, 2007 |
Current U.S.
Class: |
415/156 |
Current CPC
Class: |
F16K 7/123 20130101;
F01D 17/143 20130101; F01D 5/022 20130101; F02C 3/073 20130101;
F05D 2300/501 20130101; F16K 7/10 20130101; F02K 3/068 20130101;
F02C 3/14 20130101; F01D 17/14 20130101; F02C 9/18 20130101; F02C
6/08 20130101; F04D 27/0215 20130101; F04D 27/023 20130101; F16K
31/1266 20130101; F01D 17/141 20130101 |
Class at
Publication: |
415/156 |
International
Class: |
F03B 1/04 20060101
F03B001/04 |
Goverment Interests
[0001] This invention was conceived in performance of U.S. Air
Force contract F33657-03-C-2044. The government may have rights in
this invention.
Claims
1. A compressor for a turbine engine comprising: a compressor case;
and a bleed valve having an inlet leading from the compressor case,
the bleed valve further including a passageway leading from the
inlet and an inflatable valve selectively obstructing the
passageway.
2. The compressor of claim 1 wherein the inflatable valve
selectively obstructs the passageway continuously between a fully
open position and a fully closed position.
3. The compressor of claim I wherein the inflatable valve increases
its obstruction of the passageway as its inflation is
increased.
4. The compressor of claim 1 further including a plurality of
stages of compressor blades and a plurality of stages of compressor
vanes within the compressor case, wherein the inlet is located
between one of the stages of compressor blades and one of the
stages of compressor vanes.
5. The compressor of claim 1 wherein the bleed valve is radially
outward of the compressor case.
6. A turbine engine including the compressor of claim 1, the
turbine engine further including a fan having a plurality of fan
blades, wherein at least one of the fan blades defines a compressor
chamber extending radially therein, the compressor compressing core
airflow, wherein at least some of the core airflow from the
compressor that does not enter the inlet of the bleed valve is sent
to the compressor chamber.
7. The turbine engine of claim 6 further including an actuation
valve selectively supplying high pressure core airflow into the
inflatable valve from an area after the compressor chamber to
selectively open and close the inflatable valve.
8. A turbine engine including the compressor of claim I wherein the
bleed valve is radially inward of a bypass airflow path through the
turbine engine.
9. A bleed valve for a compressor for a turbine engine comprising:
a first member at least partially defining a passageway from an
inlet; and a valve member adjacent a portion of the passageway,
wherein the valve member is selectively moved into the passageway
to selectively obstruct the passageway upon the introduction of a
pressurized fluid to the bleed valve.
10. The bleed valve of claim 9 wherein the first member is an
annular first member and wherein the valve member is an expandable
member, the first member and the expandable member defining an
inflatable interior therebetween.
11. The bleed valve of claim 9 wherein the valve member selectively
obstructs the passageway in a continuously manner between a fully
open position and a fully closed position.
12. The bleed valve of claim 9 wherein the valve member increases
its obstruction of the passageway as the pressure of the
pressurized fluid is increased.
13. A compressor for a turbine engine including the bleed valve of
claim 9, wherein the compressor includes a plurality of stages of
compressor blades and a plurality of stages of compressor vanes,
wherein the inlet of the bleed valve is located between one of the
stages of compressor blades and one of the stages of compressor
vanes.
14. The compressor of claim 13 wherein the bleed valve is radially
outward of the compressor blades and the compressor vanes.
15. The compressor of claim 13 wherein the valve member is an
expandable member and where the expandable member is inflated by
the pressurized fluid to selectively obstruct the passageway.
16. A method for controlling bleed air from a compressor of a
turbine engine including the steps of: a) supplying a fluid to an
inflatable member adjacent a bleed air passageway that leads from
an interior of the compressor; and b) controlling a pressure of the
fluid within the inflatable member to selectively contract and
expand the inflatable member to selectively obstruct bleed air
through the passageway.
17. The method of claim 16 further including the step of
selectively expanding the inflatable member into the passageway to
obstruct bleed air through the passageway.
18. The method of claim 16 further including the step of varying
the pressure within the inflatable member such that the inflatable
member is continuously adjustable between a fully contracted
position in which the passageway is completely unobstructed by the
inflatable member and a fully expanded position in which the
passageway is completely obstructed by the inflatable member.
19. The method of claim 16 wherein said step b) further includes
the step of increasing the pressure within the inflatable member to
increase obstruction of the bleed air through the passageway.
20. The method of claim 16 further including the step of tapping a
supply of high pressure fluid from an area downstream of the
compressor in order to supply the fluid in said step a).
21. The method of claim 16 wherein at least a portion of the
inflatable member is radially inward of a bypass air flow path of
turbine engine.
Description
BACKGROUND OF THE INVENTION
[0002] The present invention relates to turbine engines, and more
particularly to an inflatable bleed valve for a low pressure
compressor for a turbine engine, such as a tip turbine engine.
[0003] An aircraft gas turbine engine of the conventional turbofan
type generally includes a forward bypass fan, a low pressure
compressor, a middle core engine, and an aft low pressure turbine,
all located along a common longitudinal axis. A high pressure
compressor and a high pressure turbine of the core engine are
interconnected by a high pressure shaft. The high pressure
compressor is rotatably driven to compress air entering the core
engine to a relatively high pressure. This high pressure air is
then mixed with fuel in a combustor, where it is ignited to form a
high energy gas stream. The gas stream flows axially aft to
rotatably drive the high pressure turbine, which rotatably drives
the high pressure compressor via the high pressure shaft. The gas
stream leaving the high pressure turbine is expanded through the
low pressure turbine, which rotatably drives the bypass fan and low
pressure compressor via a low pressure shaft.
[0004] Although highly efficient, conventional turbofan engines
operate in an axial flow relationship. The axial flow relationship
results in a relatively complicated elongated engine structure of
considerable length relative to the engine diameter. This elongated
shape may complicate or prevent packaging of the engine into
particular applications.
[0005] A recent development in gas turbine engines is the tip
turbine engine. Tip turbine engines may include a low pressure
axial compressor directing core airflow into hollow fan blades. The
hollow fan blades operate as a centrifugal compressor when
rotating. Compressed core airflow from the hollow fan blades is
mixed with fuel in an annular combustor, where it is ignited to
form a high energy gas stream which drives the turbine that is
integrated onto the tips of the hollow bypass fan blades for
rotation therewith as generally disclosed in U.S. Patent
Application Publication Nos.: 20030192303; 20030192304; and
20040025490. The tip turbine engine provides a thrust-to-weight
ratio equivalent to or greater than conventional turbofan engines
of the same class, but within a package of significantly shorter
length.
[0006] The compressors for turbine engines are designed at the
maximum power point. When operating at partial power points it
sometimes becomes necessary to bleed air form the back of the
compressor for stage matching reasons. At times, the rear
compressor stages cannot handle the amount of flow that the front
stages are pumping. To match flow, some air is bled off to reduce
the flow entering the rear stages. Turbine engines may also use
bleed air internally for accessory functions. Some bleed air may be
discharged radially out through some of the turbine blades or
stators for cooling purposes.
[0007] The compressor of a conventional turbine engine includes a
bleed valve assembly including a rotating and translating ring with
linkages. A large hydraulic actuator is disposed immediately
proximate the bleed valve for selectively opening and closing the
bleed valve. These bleed valve assemblies are large, heavy and
complex. Moreover, these bleed valve assemblies are not easily
packaged into the low pressure axial compressors for tip turbine
engines. Conventional bleed valves like this are also radially
inward of the bypass flow; however, the low compressor in
conventional engines dips radially inward at the aft end of
providing the room needed for the bleed valve. This is not true on
the tip turbine engine.
SUMMARY OF THE INVENTION
[0008] In a turbine engine according to the present invention, a
compressor for a turbine engine includes an inflatable bleed valve
that selectively bleeds core airflow from the compressor. The bleed
valve has an inlet leading from the compressor and a passageway
leading from the inlet. An inflatable valve includes an expandable
member that selectively obstructs the passageway based upon a
controlled supply of high pressure air to the inflatable valve. The
supply of high pressure air may be compressed core airflow from an
area downstream of the inlet to the bleed valve.
[0009] In a tip turbine engine, the inflatable bleed valve may be
located radially inwardly of the bypass airflow. The inflatable
bleed valve is small enough to fit within the cavity defined by the
splitter and the compressor case in a tip turbine engine. Although
the inflatable bleed valve is particularly beneficial for a tip
turbine engine, it could also be used in conventional turbine
engines.
BRIEF DESCRIPTION OF THE DRAWINGS
[0010] Other advantages of the present invention can be understood
by reference to the following detailed description when considered
in connection with the accompanying drawings wherein:
[0011] FIG. 1 is a partial sectional perspective view of a tip
turbine engine.
[0012] FIG. 2 is a longitudinal sectional view of the tip turbine
engine of FIG. 1 along an engine centerline.
[0013] FIG. 3 is an enlarged view of the inflatable bleed valve of
FIG. 2.
[0014] FIG. 4 is a view, similar to that of FIG. 3, of an
alterative inflatable bleed valve.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0015] FIG. 1 illustrates a general perspective partial sectional
view of a tip turbine engine (TTE) type gas turbine engine 10. The
engine 10 includes an outer nacelle 12, a rotationally fixed static
outer support structure 14 and a rotationally fixed static inner
support structure 16. A plurality of fan inlet guide vanes 18 are
mounted between the static outer support structure 14 and the
static inner support structure 16. Each inlet guide vane preferably
includes a variable trailing edge 18A.
[0016] A nosecone 20 is preferably located along the engine
centerline A to improve airflow into an axial compressor 22, which
is mounted about the engine centerline A behind the nosecone
20.
[0017] A fan-turbine rotor assembly 24 is mounted for rotation
about the engine centerline A aft of the axial compressor 22. The
fan-turbine rotor assembly 24 includes a plurality of hollow fan
blades 28 to provide internal, centrifugal compression of the
compressed airflow from the axial compressor 22 for distribution to
an annular combustor 30 located within the rotationally fixed
static outer support structure 14.
[0018] A turbine 32 includes a plurality of tip turbine blades 34
(two stages shown) which rotatably drive the hollow fan blades 28
relative a plurality of tip turbine stators 36 which extend
radially inwardly from the rotationally fixed static outer support
structure 14. The annular combustor 30 is disposed axially forward
of the turbine 32 and communicates with the turbine 32.
[0019] Referring to FIG. 2, the rotationally fixed static inner
support structure 16 includes a splitter 40, a static inner support
housing 42 and a static outer support housing 44 located coaxial to
said engine centerline A.
[0020] The axial compressor 22 includes the axial compressor rotor
46, which is mounted for rotation upon the static inner support
housing 42 through an aft bearing assembly 47 and a forward bearing
assembly 48. A plurality of compressor blades 52a-c extend radially
outwardly from the axial compressor rotor 46 within a fixed
compressor case 50. A plurality of compressor vanes 54a-c extend
radially inwardly from the compressor case 50 between stages of the
compressor blades 52a-c. The compressor blades 52a-c and compressor
vanes 54a-c are arranged circumferentially about the axial
compressor rotor 46 in stages (three stages of compressor blades
52a-c and compressor vanes 54a-c are shown in this example).
[0021] A bleed valve 57 mounted between the compressor case 50 and
the splitter 40 has an inlet 58 through the compressor case 50
between the last compressor vanes 54c and the last compressor
blades 52c. The bleed valve 57 includes an outlet 60 between the
compressor case 50 and the splitter 40. The bleed valve 57
selectively bleeds air out from the axial compressor 22 to control
the amount of compressed core airflow into the hollow fan blades
28, depending upon the requirements of the tip turbine engine 10 at
the time. A valve 61 obtains high pressure air from a conduit 62
leading from the combustor 30 and selectively supplies the high
pressure air to the bleed valve 57 to controllably close the bleed
valve 57 a selected amount. The valve 61 also selectively releases
air from the bleed valve 57 through an outlet 63 into the cavity
between the compressor case 50 and splitter 40 to selectively open
the bleed valve 57 a selected amount. Air flowing through the bleed
valve 57 from the axial compressor 22 is released in the cavity
between the compressor case 50 and the splitter 40, where it may
pass through the inlet guide vane 18 and discharge at an outer
diameter of the nacelle 12. The valve 61 could be mounted in a
variety of locations and connected via conduit to the bleed valve
57. For example, the valve 61 could be located in the nacelle 12
adjacent the combustor 30.
[0022] The fan-turbine rotor assembly 24 includes a fan hub 64 that
supports a plurality of the hollow fan blades 28. Each fan blade 28
includes an inducer section 66, a hollow fan blade section 72 and a
diffuser section 74. The inducer section 66 receives airflow from
the axial compressor 22 generally parallel to the engine centerline
A and turns the airflow from an axial airflow direction toward a
radial airflow direction. The airflow is radially communicated
through a core airflow passage 80 within the fan blade section 72
where the airflow is centrifugally compressed. From the core
airflow passage 80, the airflow is diffused and turned once again
by the diffuser section 74 toward an axial airflow direction toward
the annular combustor 30. Preferably, the airflow is diffused
axially forward in the engine 10, however, the airflow may
alternatively be communicated in another direction.
[0023] The tip turbine engine 10 may optionally include a gearbox
assembly 90 aft of the fan-turbine rotor assembly 24, such that the
fan-turbine rotor assembly 24 rotatably drives the axial compressor
22 via the gearbox assembly 90. In the embodiment shown, the
gearbox assembly 90 provides a speed increase at a 3.34-to-one
ratio. The gearbox assembly 90 may be an epicyclic gearbox, such as
a planetary gearbox as shown, that is mounted for rotation between
the static inner support housing 42 and the static outer support
housing 44. The gearbox assembly 90 includes a sun gear 92, which
rotates the axial compressor rotor 46, and a planet carrier 94,
which rotates with the fan-turbine rotor assembly 24. A plurality
of planet gears 93 each engage the sun gear 92 and a rotationally
fixed ring gear 95. The planet gears 93 are mounted to the planet
carrier 94. The gearbox assembly 90 is mounted for rotation between
the sun gear 92 and the static outer support housing 44 through a
gearbox forward bearing 96 and a gearbox rear bearing 98. The
gearbox assembly 90 may alternatively, or additionally, reverse the
direction of rotation and/or may provide a decrease in rotation
speed.
[0024] A plurality of exit guide vanes 108 are located between the
static outer support housing 44 and the rotationally fixed exhaust
case 106 to guide the combined airflow out of the engine 10 and
provide forward thrust. An exhaust mixer 110 mixes the airflow from
the turbine blades 34 with the bypass airflow through the fan
blades 28.
[0025] The bleed valve 57 is shown in more detail in FIG. 3. The
bleed valve 57 includes a passageway 112 between the inlet 58 and
the outlet 60. In this embodiment, the passageway 112 extends
generally axially forward, such that the inlet 58 is located aft of
the outlet 60, however, alternative orientations could be used. An
opening 114 is formed on the outer diameter of the passageway 112
and an inflatable, annular valve 116 is mounted over the opening
114. The valve 116 includes a rigid outer annular ring 118 to which
is mounted a seal 120. A flexible, expandable ring 122, radially
inward of the seal 120, defines an inflatable interior 124 between
the ring 122 and the seal 120. The valve 61 (FIG. 2) selectively
supplies high pressure air to the interior 124, thereby selectively
causing the ring 122 to expand through the opening 114 and obstruct
the passageway 112 by a controlled amount. The ring 122 can
selectively be expanded any amount between an uninflated, fully
retracted position, as shown, and a fully expanded, filly inflated
position where the passageway 112 is completely closed. Air flowing
through the bleed valve 57 from the axial compressor 22 is released
in the cavity between the compressor case 50 and the splitter 40 or
may be used for accessory functions, thereby reducing the amount of
core airflow into the inducer 66 and the hollow fan blades 28.
[0026] FIG. 4 shows a bleed valve 157 according to a second
embodiment of the present invention, which could also be used in
the tip turbine engine 10 of FIGS. 1-2. In this embodiment, the
passageway 212 of the bleed valve 157 extends radially outwardly,
such that the outlet 160 is substantially radially aligned with the
inlet 158. The flexible ring 222 is similarly selectively
expandable to control the amount of core airflow bled from the
axial compressor 22. In FIG. 4, the flexible ring 222 is shown in
the uninflated, open position as reference numeral 222 and in the
inflated, closed position as 222'. Again, it is noted that the
flexible ring 222 is also selectively adjustable to any point
between fully open and filly closed.
[0027] Referring to FIG. 2, in operation, core airflow enters the
axial compressor 22, where it is compressed by the compressor
blades 52a-c. To control the core airflow into the combustor 30,
the bleed valve 57 (or, optionally bleed valve 157 from FIG. 4) is
selectively opened or closed a selected amount. Bleed air is
discharged through the inlet guide vane 18 and/or may be used for
accessory functions. The compressed air from the axial compressor
22 that is not bled off enters the inducer section 66 in a
direction generally parallel to the engine centerline A, and is
then turned by the inducer section 66 radially outwardly through
the core airflow passage 80 of the hollow fan blades 28. The
airflow is further compressed centrifugally in the hollow fan
blades 28 by rotation of the hollow fan blades 28. From the core
airflow passage 80, the airflow is turned and diffused axially
forward in the engine 10 by the diffuser section 74 into the
annular combustor 30. The compressed core airflow from the hollow
fan blades 28 is mixed with fuel in the annular combustor 30 and
ignited to form a high-energy gas stream.
[0028] The high-energy gas stream is expanded over the plurality of
tip turbine blades 34 mounted about the outer periphery of the
fan-turbine rotor assembly 24 to drive the fan-turbine rotor
assembly 24, which in turn rotatably drives the axial compressor 22
either directly or via the optional gearbox assembly 90. The
fan-turbine rotor assembly 24 discharges fan bypass air axially aft
to merge with the core airflow from the turbine 32 in the exhaust
case 106.
[0029] In accordance with the provisions of the patent statutes and
jurisprudence, exemplary configurations described above are
considered to represent a preferred embodiment of the invention.
However, it should be noted that the invention can be practiced
otherwise than as specifically illustrated and described without
departing from its spirit or scope.
* * * * *